The invention relates generally to hot gas path components for turbine assemblies and, more particularly, to synergistic approaches to cool the hot gas path components.
Exemplary gas turbine engines are used for aircraft or stationary power applications, and engine efficiency is a key design criteria for both applications. The efficiency of gas turbine engines improves with increased temperature of the combustion gas flow. However, a limiting factor in the gas flow temperature is the high temperature capability of the various hot gas path components, such as the turbine stator and rotor airfoils. Stator airfoils are also known as vanes or nozzles, rotor airfoils are also known as blades or buckets.
Various approaches to cooling the hot gas path components have been proposed and implemented to increase the upper operating temperature of the engines. Several of these approaches are reviewed in commonly assigned U.S. Pat. No. 5,690,472, Lee, “Internal Cooling of Turbine Airfoil Wall Using Mesh Cooling Arrangement.” These cooling techniques typically involve bleeding compressed air off the compressor to serve as a coolant. However, by bypassing the compressed air around the engine's combustion zone to cool the hot gas path components, the overall efficiency of the engine is reduced. Accordingly, it is desirable to increase the cooling effectiveness of the hot gas path components, in order to improve overall engine efficiency.
One beneficial cooling technique is mesh cooling, as described, for example in U.S. Pat. No. 5,690,472, which is cited above, and in U.S. Pat. No. 5,370,499, Lee, “Film Cooling of Turbine Airfoil Wall using Mesh Cooling Hole Arrangement.” However, a need for additional improvement in cooling of hot gas path components remains. This need is especially strong for cooling thin airfoil walls and/or regions of limited accessibility, such as the trailing edges of airfoils. Accordingly, it would be desirable to provide enhanced cooling effectiveness for hot gas components.